This invention relates to coatings capable of use on components exposed to high temperatures, such as the hostile thermal environment of a gas turbine engine. More particularly, this invention is directed to a thermal barrier coating (TBC) capable of exhibiting resistance to thermal cycling and infiltration by contaminants, for example, of types that may be present in the operating environment of a gas turbine engine.
The use of thermal barrier coatings (TBCs) on components such as combustors, high pressure turbine (HPT) blades, vanes and shrouds is increasing in commercial as well as military gas turbine engines. The thermal insulation provided by a TBC enables such components to survive higher operating temperatures, increases component durability, and improves engine reliability. TBCs are typically formed of a ceramic material and deposited on an environmentally-protective bond coat to form what is termed a TBC system. Bond coat materials widely used in TBC systems include oxidation-resistant overlay coatings such as MCrAlX (where M is iron, cobalt and/or nickel, and X is yttrium or another rare earth element), and diffusion coatings such as diffusion aluminides that contain aluminum intermetallics. Bond coat materials are typically selected to be capable of forming a continuous and adherent oxide scale on their surface to promote the adhesion of the ceramic coating to the bond coat. The oxide scale can be formed by subjecting the bond coat to an oxidizing environment, such that the scale is sometimes referred to as a thermally-grown oxide (TGO).
Notable examples of ceramic materials for TBCs include zirconia partially or fully stabilized with yttria (yttrium oxide; Y2O3) or another oxide, such as magnesia, ceria, scandia and/or calcia, and optionally other oxides to reduce thermal conductivity. Binary yttria-stabilized zirconia (YSZ) is widely used as a TBC material because of its high temperature capability, low thermal conductivity, and relative ease of deposition. Zirconia is stabilized to inhibit a tetragonal to monoclinic crystal phase transformation at about 1000° C., which results in a volume change that can cause spallation. At room temperature, the more stable tetragonal phase is obtained and the monoclinic phase is minimized if zirconia is stabilized by at least about six weight percent yttria. A stabilizer (e.g., yttria) content of seventeen weight percent or more ensures a fully stable cubic crystal phase. The conventional practice has been to partially stabilize zirconia with six to eight weight percent yttria (6-8% YSZ) to obtain a TBC that is adherent and spallation-resistant when subjected to high temperature thermal cycling. Furthermore, partially stabilized YSZ (e.g., 6-8% YSZ) is known to be more erosion-resistant than fully stabilized YSZ (e.g., 20% YSZ).
Various process can be used to deposit TBC materials, including thermal spray processes such as air plasma spraying (APS), vacuum plasma spraying (VPS), low pressure plasma spraying (LPPS), and high velocity oxy-fuel (HVOF). TBCs employed in the highest temperature regions of gas turbine engines are often deposited by a physical vapor deposition (PVD), and particularly electron beam physical vapor deposition (EBPVD), which yields a columnar, strain-tolerant grain structure that is able to expand and contract without causing damaging stresses that lead to spallation. Similar columnar microstructures can be produced using other atomic and molecular vapor processes, such as sputtering (e.g., high and low pressure, standard or collimated plume), ion plasma/cathodic arc deposition, and all forms of melting and evaporation deposition processes (e.g., laser melting, etc.). TBCs formed by the various methods noted above generally have a lower thermal conductivity than a dense ceramic of the same composition as a result of the presence of microstructural defects and pores at and between grain boundaries of the TBC microstructure.
Under service conditions, hot section engine components protected by a TBC system can be susceptible to various modes of damage, including erosion, oxidation and corrosion from exposure to the gaseous products of combustion, foreign object damage (FOD), and attack from environmental contaminants. The source of environmental contaminants is ambient air, which is drawn in by the engine for cooling and combustion. The type of environmental contaminants in ambient air will vary from location to location, but can be of a concern to aircraft as their purpose is to move from location to location. Environmental contaminants that can be present in the air include sand, dirt, volcanic ash, sulfur in the form of sulfur dioxide, fly ash, particles of cement, runway dust, and other pollutants that may be expelled into the atmosphere, such as metallic particulates, for example, magnesium, calcium, aluminum, silicon, chromium, nickel, iron, barium, titanium, alkali metals and compounds thereof, including oxides, carbonates, phosphates, salts and mixtures thereof. These environmental contaminants are in addition to the corrosive and oxidative contaminants that result from the combustion of fuel. However, all of these contaminants can adhere to the surfaces of the hot section components, including those that are protected with a TBC system.
In order for a TBC to remain effective throughout the planned life cycle of the component it protects, it is important that the TBC has and maintains integrity throughout the life of the component, including when exposed to contaminants. Some contaminants may result in TBC loss over the life of the components. For example, particulates of calcia (CaO), magnesia (MgO), alumina (aluminum oxide; Al2O3) and silica (silicon dioxide; SiO2) are often present in environments containing fine sand and/or dust. When present together at elevated temperatures, calcia, magnesia, alumina and silica can form a eutectic compound referred to herein as CMAS. A particular composition that has been identified for CMAS contains about 35 mol % CaO, about 10 mol % MgO, about 7 mol % Al2O3, and about 48 mol % SiO2, along with about 3 mol % Fe2O3 and about 1.5 mol % NiO. CMAS has a relatively low melting temperature, such that during turbine operation the CMAS that deposits on a component surface can melt, particularly if surface temperatures exceed about 2240° F. (1227° C.). Molten CMAS is capable of infiltrating the porosity within TBCs. For example, CMAS is capable of infiltrating into TBCs having columnar structures, dense vertically-cracked TBCs, and the horizontal splat boundaries of TBCs deposited by thermal and plasma spraying. The molten CMAS resolidifies within cooler subsurface regions of the TBC, where it interferes with the compliance of the TBC and can lead to spallation and degradation of the TBC, particularly during thermal cycling as a result of interfering with the ability of the TBC to expand and contract. In addition to loss of compliance, deleterious chemical reactions with yttria and zirconia within the TBC, as well as with the thermally-grown oxide at the bond coating/TBC interface, can occur and cause degradation of the TBC system. Once the passive thermal barrier protection provided by the TBC has been lost, continued operation of the engine will lead to oxidation of the base metal beneath the TBC system, which may ultimate lead to failure of the component by burn through cracking.
Attempts to mitigate the effect of the CMAS on high pressure turbine blades and shrouds have included the application of a thin layer of alumina on the surface of the TBC to increase the melting point of CMAS by about 100 to 150° F. (38° C. to 66° C.), for example, as reported in U.S. Pat. No. 5,660,885. The addition of the alumina layer provides an increase in operating temperature of up to about 2400° F. (1316° C.) with reduced infiltration of liquid CMAS. However, grinding during manufacture and assembly, as well as grinding and rubbing with turbine shrouds during gas turbine engine operation, result in the use and reliance on the alumina layer difficult and impractical. In addition, the alumina layer adds manufacturing cost and complexity, especially for turbine blades that are subjected to gas and particle erosion and may have different requirements for the alumina coating in order to minimize erosion. In addition, thicker alumina layers are subject to coefficient of thermal expansion mismatches within the TBC coating system, resulting in thermal strains during cycling.
In view of the above, it can be appreciated that there are certain problems, shortcomings or disadvantages associated with the prior art, and that it would be desirable if systems and methods were available that are capable of promoting the resistance of components to contaminants, such as CMAS, and particularly gas turbine engine components that operate at temperatures above the melting temperatures of contaminants.